Referring to FIG. 1, a gas turbine engine is generally indicated at 10 and comprises, in axial flow series, an air intake 11, a propulsive fan 12, an intermediate pressure compressor 13, a high pressure compressor 14, a combustor 15, a turbine arrangement comprising a high pressure turbine 16, an intermediate pressure turbine 17 and a low pressure turbine 18, and an exhaust nozzle 19.
The gas turbine engine 10 operates in a conventional manner so that air entering the intake 11 is accelerated by the fan 12 which produce two air flows: a first air flow into the intermediate pressure compressor 13 (and engine core), and a second air flow to the bypass stream, which provides propulsive thrust. The intermediate pressure compressor compresses the air flow directed into it before delivering that air to the high pressure compressor 14 where further compression takes place.
The compressed air exhausted from the high pressure compressor 14 is directed into the combustor 15 where it is mixed with fuel and the mixture combusted. The resultant hot combustion products then expand through, and thereby drive, the high, intermediate and low pressure turbines 16, 17 and 18 before being exhausted through the nozzle 19 to provide additional propulsive thrust. The high, intermediate and low pressure turbines 16, 17 and 18 respectively drive the high and intermediate pressure compressors 14 and 13 and the fan 12 by suitable interconnecting shafts.
It will be understood from the above that achieving most efficient engine operation is a prime objective in both the compressors and turbines. Stator vanes are utilised in order to re-orientate and direct fluid flow for most efficient operation in the engine. These stator vanes may be shrouded and either take the form of single vanes or cast spans of several vanes or single vanes welded or brazed into spans. In either event, each vane or span of vanes at its mounting end is formed upon a platform section, and in creating the vane structure about the circumference of the engine, there are a number of gaps between these platform sections. These gaps are provided to account for thermal expansion and vibration damping in operation of the engine. FIG. 2 shows inter-platform gaps 100 between platforms (101) upon which vanes 103 are presented.
Unfortunately, such gaps 100 also create a leakage path connecting the flow through the blade passage to the shroud ring cavity (22). FIG. 3 illustrates such leakage as broken lines 21 which pass into the cavity 22 via the gap created between a platform 23 of a first blade 24 and the next adjacent platform (not shown). As can be seen the leakage path creates the flow 21 into the cavity 22 and air from the cavity 22 then re-enters a vane passage 25. In such circumstances, the leakage flow 21 increases aerodynamic losses within each vane 24 row and so reduces compressor efficiency within an engine.
Existing shroud vane designs do not seal the blade passage flow 25 from the shroud ring cavity 22 and so are subject to loss generating leakage flow through the inter vane gap described above. This leakage flow will reduce compressor efficiency. It should be understood with regard to aircraft operations, additional weight in the form of sealing components may require consideration as to whether the sealing benefits are justified. As indicated above there is a necessity for platform gaps in order to accommodate for thermal expansion and for vibration damping, so that consequently the inherent aerodynamic losses may be simply accepted as unavoidable.
As also illustrated in FIG. 3, it will be noted that anti-fret liners 26 are provided between mounting portions 27 of the mounting shroud 28 and reciprocal mounting portions 29 of the platform 23. These anti-fret liners act to inhibit vibration or otherwise induced abrasive degradation in the mounting association between the members 27, 29.